Auxilliary carriage arrangement for a missile



W. H. GOSS Aug. 18, 1959 AUXILLIARY CARRIAGE ARRANGEMENT FOR A MISSILE Original Filed April 3, .1946

3 Sheets-Sheet vl mau/3H. Goss @Sygl I' NALI'TORNEYv W. H. Goss Aug. 18, 1959 AUXILLIARY CARRIAGE ARRAGEMENT lFOR A MISSILE Original Filed April 3, 1946 C5 Sheets-Sheet 2 m m a m, m H@ R A U m i@ n@ W. H. GOSS Aug. 18, 1959 AUXILLIARYxCARRIAGE ARRANGEMENT FOR A MISSILE Original Filed April 3, 1946 3 Sheets-Sheet 3 United States Patent AUXILIARY CARRIAGE ARRANGEMENT FOR'A MISSlLE S Claims. (Cl. 102-49) The. present.inventionrelates generally toaerial mis-V siles and more particularly to an improved ramjet missile and method of sustaining combustion therein. y

Thisapplication is divided from United States patent applicationiSerial Number 659,188, led April 3, 1946, by. David H. Sloan and Wilbur H. Goss for Ram let Projectile.

Ramjet missiles developed up to -the present time have been incapable of overcoming drag to the extent necessary. to develop thrust usable for permitting climbing. Moreover, such. missiles involve the use of a plurality of moving partsand, without exception, they have been incapablelofdevelopingfeven' near-maximum power from the'.l fuel: einployed,.with thevv overall result thaty up to the .presenttime*asuccessful practical ramjet missile hasnot been available.

One ofthe objects ofthe present invention, therefore, is to provide' a ramjet missile (athodyd) which will be light in weightand whichzwill be capable of delivering large enough force to overcome drag and permit climbing of said missile.

Another object of the invention is to provide a ramjet missilehaving no moving'parts.

Still another object resides in the provision of a missile offthis type which' will obtain a higherspeciiic impulse from the fuel` employed than may be obtained from a` rocket.

Asa further object, the invention provides vimproved fuellstorage and feeding means which will make use of air pressure present during ight of the craft through the atmosphere, with the result that smooth, uniform fuel feeding will be assured.

As a further object, the invention contemplates a ramjet-missile capable of obtaining its supply of oxygen, for supporting the burning of the fuel in ight, from the atmosphere, with the result that added space for carrying a greater load or additional fuely will be made avail'- able.

Other objects and many of the attendant advantagesy of thisv invention will be readily appreciated as the same becomes better understood by reference lto the following detailed'description when considered inconnectionwith the accompanying drawings, wherein:

Fig. 1 is a' diagrammatic View of the improved'ramjet missile;

Fig. 2 is a side elevation illustrating the missile assembly;

Figs. 3 and 3A comprise a longitudinal sectional View of one embodiment ofthe invention, partly schematic, having 'a' two-stage'igniter shown partly in elevation, the views being connectable on the lines A-A and B'-B;

Fig. 4 is a fragmentary perspective view of'the'invention, showing particularlythe mountingfor thev two-stage igniter;-

Fig; 5 is a longitudinal sectional view of a modified embodiment of the invention, the View being broken away at its midand rear end portionsand the nose or windshield being shown fragmentarily;

'2,899,898 Patented Aug. 18, 1 959 of theY launching. ramp being shown fr'agmentarily;l and Figs. 8 and 9 are perspective and endviews, respectively, showing ay missile in operative position in a car riage, portions of Fig. 8 beingshown in section.

Referring now more particularly to thedrawings and first to Fig.` l thereof, wherein the ramjet missile is shown. schematically, let it be assumed that the missileis moving through the atmosphere. at supersonicvelocity W inthe direction `shown by the farrow. Since the velocity; atA` is supersonic,` ta shock wave will probablyform just in front of the intake (at the forward end ofthe diiuser to beV described presently) of the missile. If the shock wave isplane, kthe velocity of the air atl 1-will beisubsonic but pressure P1,- density i1-andk temperature T1 of the air will kbe increased. The velocity of the air is next further slowed by the diifuseruntil it reaches the desired burning velocity atB. Fuel `is injected betweenB andl B and-.is burned between B and C. The burning should take place atconstant pressure; From C the hot' exhaust gases expand through a constriction or nozzle E andtheir velocity is kincreased by said nozzle. Broadly speaking; the thrust whichl is developed bythe high velocity gas/W arises from the fact that the momentum of these gases -is' greater at vD than at the intake tothe diffuser.

In the embodiment-ofthe improvedmissile as shownin-Figs. 2, 3 and-'4Y of the drawings, the numeral 1 0/inV dicatesv ar tubular body which is formed of heavy'ga'uge sheet metal and which is open at bothl ends. The body may be of variousselected lengths and diameters, those of asuccessful embodimentzof'the invention havingbeen aboutten feet long and six-inches in diameter. Thereaf. end portion' ofzthe body Itlis reduced andthenslightly flaredvtodeiine a'nozzle 11. Fins 12, preferably'fourin. number, are mounted on the rear endiportionof the-bodyatanglesv ofninetyV degrees and are for the `purpose of assuring that'the missilewill be maintained on its course'v in flight. `The iins12 are generally rectangularshapeddri side elevation and'may have anysuitable dimensions Partially closing the forward end of the body 10Yis -a nose' assembly 13 which includes ya' nose ory ogive 1'4-0'f plastic, metal. or other suitable materiaLand-a metallange 1S Vfor attaching said ogive lto the body 10? The flange 15'4 is formed with right angular boresr1`6' which communicate between the outer annulary surface of the flange. andthe annular rear` wall thereof. The bores are preferably-four in number, spacedat ninety degree in-' tervals about the flange. Y Preferably, four Pit'ottubes 1'7' are' connected with' the vflange 1S and are mounted with* their outer endsl direc/tedl forwardly and vtlieirrearl or inner. ends soldered or otherwise secured to the flange in:V surrounding relation to the mouths ofthe bores 16. The purpose" of-the' bores 16 andflPitot tubes 17' will E be "def scribed 'in' more detail hereinafter.-

The ogive or nose 14 is Vformed with a central passage 18 constricted at its mid portion and having its rear end' communicating with a generally frusto-conical diiuser 19 which has its forward end anchored in a central open-A ing in the liange 15Aand'which' extends throughout substantially one-third the length off'the body 1'0. As will be seen; the open rear end of the diffuser terminates in close vspacedY relation to the inner wall of theV body'lliE and is-held in rigid position by a mounting ring'20.`

The'fuel 'system for the-improved missile includes an annular@ tankv 21 `located immediately rearwardly off the noseassemlialyv Band having its inner wall? delined"by"'aportioirr'of'thediffuser 19. The tank includes a tubular outer wall 22 ofrubben or the' like, which surrounds the diffuser in spaced relation thereto and which has its rear end margin clamped by a pair of rings 23 and 24, the ring 24 being anchored to the wall of the diffuser 19. The forward end margin of the tank 22 is clamped about the rear end portion of the flange 15 by a clamping ring 25 which is formed with longitudinal ports 26 therein, said ports being arranged in spaced relation about the ring 25. An annular series of jets 27 is mountedin the forward end of the diffuser 19 and each jet communicates with the interior of the tank 21 at its forward end. It will now be understood that as the missile moves through the air in flight, air flow will enter the Pitot tubes 17 and pass through the bores 16 and ports 26 into the annular space between the wall 22 of the tank and the wall 10, with the result that the wall 22 will be compressed inwardly for forcing fuel such asV gasoline in the tank 21 through the nozzles or jets 27, where it will be mixed with air entering the central opening 18 andforced rearwardly in the diffuser 19, in a manner and for a purpose to be discussed in more detail hereinafter.

The fuel ignition system employed with this embodiment of the invention will now be briefly described. In this connection, it should be understood that various types and methods of fuel ignition may be employed in lieu of the system embodied in this description. The fuel ignition system illustrated includes a duct 28 which extends downstream in the body 10 concentrically with respect thereto from a point in relatively closely spaced relation to the rear end of the diffuser 19. At its rear end portion, the duct 28 is supported in body 10 by streamlined supports 29 (Fig. 4) arranged at angles of ninety degrees with respect to each other. The duct 28, as shown, extends substantially two-fifths the distance from the rear end of the diffuser toward the rear end of the body. Shorter or longer ducts may be employed if desired. Positioned within the duct 28, preferably by four streamlined inner supports 30, is an igniter holder 31 which is of tubular shape and which extends substantially one-third the length of the duct. The holder 31 is closed at its forward end by an ogive-shaped nose 32 which extends partly into the diffuser 19 substantially axially thereof. An igniter 33 is carried within the holder 31 and is preferably of a construction similar to tracers commonly used in missiles. The igniter 33 is, however, electrically detonated and, for the purpose of effecting electrical detonation, a pair of wires 34, connected with a primer (not shown) in the igniter, extend rearwardly therefrom for connection in a circuit including a battery 35 and a switch 36, said battery and switch, of course, being located a safe distance away from the launching ramp or platform to be described presently.

Reference is now had to Figures 7, 8 and 9 wherein there is illustrated a launching carriage for effecting launching of the improved missile. A satisfactory launching carriage is shown generally at 37 and includes a cluster of preferably four rocket motors 38 of the type disclosed in copending applications of C. C. Lauritsen, Serial Numbers 481,644 and 481,652, filed April 2, 1943, now Patent Numbers 2,464,181 and 2,458,475, respectively. The rocket motors employed up to the present time have been of the -inch type. It should be understood, however, that launching carriages employing a greater or less number of rocket motors, and rocket motors of different sizes have been and will be used as occasion demands. The rocket motors are connected at the innermost portions of their corresponding rear ends by a thrust plate 39 having apertures 40 therein, and at their corresponding forward ends by a manifold 41. The manifold 41 comprises cups 42 which close the forward ends of the rocket motors 38, and arcuate connecting pipes 43. The manifold 41 serves the twofold function of connecting the forward end portions of the rocket motors so as to assure uniform burning and of providing a support for spacing the forward ends of the motors with respect to each other and defining a pocket or recess for receiving the rear end portion of the ramjet missile 10.

When the missile is ready for launching, it is positioned in the carriage with its rear end in engagement with the thrust plate 39. If desired, Wedges W of Wood or other suitable material may be placed at spaced points about the missile near the forward end of `the carriage so as to assure that the missile Willbe positioned concentrically with respect to the surrounding motors 38.

The carriage 37, with the improved ramjet missile therein, is ordinarily discharged from a suitable ramp which includes a pair of spaced, parallel guide rails 44 engageable by bifurcatedopposed guides 45, a pair of said guides being carried on the manifold 41, and a second pair being carried on a connecting strap 46 surrounding the motors 38 near their corresponding rearends.,

To assure that the carriage 37 will be maintained on a true course in flight, a fin assembly including fins 47 is carried at the rear end of the rocket motor assembly.

To effect launching of the missile 10 from the ramp, the rocket motors 38 are ignited by remote control with suitable electrical equipment of conventional design (not shown). At the same time, the igniter 33 is ignited by closing the switch 36 from a remote position. After ignition, the rocket motors 38 will cause the carriage 37 with the missile 10 therein bearing against the thrust plate 39, to be moved at a high rate of speed along the ramp rails 44 and to be discharged from the ramp at the ends of said rails. The carriage 37 will impart flight to the missile 10 for a period of one or two seconds; that is, until the propellant in the motors 38 is completely burned. The carriage and the missile Will move through the air at such a high rate of speed that the velocity of air flow through the central opening or throat 18 will be sufficient to support uniform burning of fuel, forced into the mouth of the diffuser 19 through the jets 27 by compression of the tank 22, after the aerodynamic drag has stripped the carriage 37 from the missile. Uniform burning of fuel in the combustion chamber rearwardly ofthe diffuser 19 will provide sufficient thrust to drive the missile 10 through the air at supersonic speeds.

More particularly, as the missile travels through the air, air flow at high velocity will enter the diffuser 19 through the throat 18 and will be mixed with the fuel from jets 27 in the chamber defined between the jets 27 and the igniter 33. This mixture of air and fuel, at relatively, low velocity and high pressure, will be ignited by the igniter 33 and the products of combustion will be ejected through the nozzle 11 at sufficiently higher velocity than that of the air flow entering the throat 18 to provide the necessary thrust for driving the missile through the atmosphere at supersonic speeds.

It is desired to point out that the use of the ignition system embodying the duct 28 may assure a higher degree of burning, because a portion of the air flow from the diffuser 19 is passed directly about the igniter 33 and thus heated to a higher temperature than the mixture of air and fuel passing outside said duct 28, with the result that preheating of a portion of `the air flow is effected and a more complete and effective combustion takes place.

Referring now to the modification of the invention illustrated in Figs. 5 and 6, the numeral 50 indicates the cylindrical barrel or body of this device, the body being quite similar to the body 10 of the previously described embodiment of the invention. The body 50 is closed lat its forward end by an annular plug 51 having an axial opening 52 enlarged and slightly outwardly flared to define an annular recess and a shoulder for receiving the forward end portion of a diffuser element53. Injection nozzles 54, arranged in a circumferentially spaced series in the forward end portion of the diffuser element 53, communicate with the interior of a fuel tank the inner cylindrical wall of which is defined by a portion of the diffuser element 53. As will be seen, the nozzles 54 include generally polygonal blocks having rearwardly inclined walls 55 for protecting the jets` 56 from any explosive blasts entering the diffuser 53 from the combustion chamber. A circular plug 57, of rubber or other suitable material, closes the corresponding outer ends of the jets 56. This plug 57 is displaced and blown out of said missile by air flow passing through said missile after it has been launched into ilight.

The fuel tank of this embodiment of the invention comprises a cylindrical, preferably rubber, outer wall 58 secured at its opposite ends to reduced terminal rings 59 and 60, also of rubber. The free end margin of the ring 59 is clamped on the rear portion of the plug 51, which is reduced and tapered as shown at 61, by a clamping ring 62. As will be seen, the clamping ring 62 is formed with a series of longitudinally extending bores 63 which communicate with bores 64 in the plug 5.1. The bores 64 communicate with Pitot tubes 65 mounted in said bores and having their corresponding outer ends directed forwardly, or upstream. The rear ring 60 is clamped between confronting inclined surfaces of clamping rings 66 and 67, the ring 67 being fastened tightly to the diffuser element S3 near its rear end portion. The ring 66 is tightly fitted about the rear margin of the terminal ring 60 and, to prevent pressure leakage, the ring 66 is formed with a groove 68 in which is mounted an annular gasket or sealing ring 69 which engages the inner wall of the body 50.

In order to assure uniform ilow of fuel to the nozzles 54 regardless of the distribution of pressure about the wall 58 of the fuel tank, there is provided a plurality of yducts or conduits 70. As best seen in Fig. 6, the ducts 70 are of substantially U-shape and extend throughout the major portion of the length of the tank along its inner wall S3. The ducts 70 are each provided with a plurality of spaced inlet ports 71 (Fig. 5) at their corresponding rearward ends. To assure that the inner ends of the nozzles 54 will be protected in the event that the outer wall 58 and the ring 59 should be completely collapsed by air pressure entering the Pitot tubes and surrounding the tank, there is provided a sheet of foraminous material 72. The sheet 72 extends throughout substantially the entire length of the tank and thus provides effective means for assuring free dow of fuel to the nozzles from any point along the length of the tank.

As in the previously described embodiment of the invention, a nose or windshield is fitted in the forward end of the body 50. More specifically, a windshield 73 is fitted in an enlarged recess 74 in the forward end of the plug 51 and has a central passage 73 communicating with the diffuser 53. The operation of this embodiment of the invention is quite similar to that of the embodiment previously described. When the missile is launched, in the manner of the first described embodiment, air flow will enter the diffuser 53 from passage 73' and will blow out the plug 57, with the result that fuel from tank 58 may be fed through jets 56 to the interior of the forward end of the diffuser. This fuel will, of course, be blown rearwardly `by the air ow and mixed therewith. In view of the fact that the ducts 70 and the foraminous sheet 72 are provided, fuel ow will be uniform to the nozzles 54 with the result that, in so far as fuel ow is concerned, uniform burning will be assured.

Obviously many modifications and variations of the present invention are possible in the light of the above teachings. It is therefore to be understood that within the scope of the appended claims the invention may be practiced otherwise than as specifically described.

What is claimed is:

1. In a ramjet missile including an open-ended body, and means for sustaining combustion in the body after said missile has been launched for imparting thrust to said body for sustaining the ight thereof; detachable means on the body for launching said missile into flight, said detachable means comprising a carriage having a plurality of rocket motors arranged in a cluster, a thrust plate connecting the rocket motors at their rear ends, a manifold connecting said rocket motors at their forward ends, said thrust plate and manifold defining a receptacle for said missile, and ns mounted on the carriage for directing flight thereof.

2. In a self-propelled aerial missile having a sustaining propulsion motor; in combination, a separable launching carriage for said missile, said carriage comprising a number of auxiliary rockets, means securing said auxiliary rockets to each other at their forward ends, other means, including a thrust plate, securing them to each other at their rear ends, said auxiliary rockets being symmetrically spaced, to provide room for the aerial missile amid said auxiliary rockets, said plate abutting the rear end of said missile.

3. A self-propelled aerial missile arrangement as set forth in claim 2, wherein the separable launching means is provided with guides, designed to run along stationary guide rails parallel to the missile axis, during the initial launching period.

4. A self-propelled guided missile, comprising, a body, said body including a plurality of axially arranged chambers, means for sustaining combustion in the rear chamber of said body, airfoils mounted on said body, a separable launching carriage initially supplying thrust to -said missile, said carriage including a thrust plate abutting the rear of the body, a plurality of launching rockets secured to said carriage in spaced relation to each other and to said missile and extending forward from said plate substantially in parallelism with the axis of the missile, and manifolds connecting the forward ends of said launching rockets, said rockets and said manifolds jointly forming a ring of sufficient inside diameter to clear the airfoils.

5. A self-propelled supersonic missile, comprising, an l elongated body, said body including a plurality of sequentially acting chambers, a sustaining thrust motor in the rear chamber of said body, a separable launching carriage receiving the rear end of said body, a plurality of launching rockets secured in said carriage in symmetrical relation to each other and to said sustaining motor, a plate cornected to said rockets and abutting the rear end of said body to transmit the thrust of said launching rockets to said body, and manifolds connecting the front ends of said launching rockets.

No references cited. 

